This invention relates to a gas turbine blisk, and, more particularly, to a gas turbine blisk in which ceramic foam blades are integrally affixed to a disk.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor. In a more complex version of the gas turbine engine, the compressor and a high pressure turbine are mounted on one shaft, and a bypass fan and a low pressure turbine are mounted on a separate shaft. In any event, the hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
There are two approaches to the assembly of the blades and disk of an axial-flow compressor or axial-flow turbine. In one approach, the blades are mechanically affixed to the disk, using a dovetail attachment or other structure. In the other approach, the blades are integral with the disk. The second approach having an integral blade/disk structure, called a xe2x80x9cbliskxe2x80x9d in the art, is favored in those situations where it may be used because the mechanical attachment of the first approach may add a substantial amount of weight to the structure. Both approaches have been used in the compressor stage. For example, high-pressure titanium-alloy blisks are used in the high-pressure compressor stages of some engines.
Blisks are not conventionally used in the high-pressure or low-pressure turbine stages. The materials used in the disk and the blades must be quite different in characteristics to achieve optimal performance. There has been no approach available to join blades to disks that has been found operable in the extreme loading and environmental conditions experienced in the turbine stages. Reinforcement of the turbine blades with ceramic fibers has not been generally successful because of thermal mismatch differences between the base metal and the ceramic.
There is therefore a need for an improved approach to the fabrication of a gas turbine blisk. The present invention fulfills this need, and further provides related advantages.
The present invention provides an integral gas turbine blade/disk (xe2x80x9cbliskxe2x80x9d) and a method for its fabrication. The blades are integrally joined to the disk. The blades are made of a material which has reduced weight as compared with conventional turbine blades and may be tailored for even further weight reduction. The blade material has excellent high-temperature mechanical properties and also is impact resistant. Uncooled blades may be used in some stages of the turbine where it was previously necessary that the blades be cooled. The use of protective coatings may be unnecessary in some stages of the turbine where they were previously necessary, reducing weight and cost. The blisk of the invention is suitable for use in high-pressure and low-pressure turbine stages, as well as the compressor stages. The introduction of blisk technology into the turbine stages significantly improves gas turbine performance by reducing the weight of the turbine disk and blades. The reduction in weight in the turbine disk and blades also results in reduced bearing weight and structural support weight.
A blisk comprises a disk and at least one blade integrally affixed to the disk. Each blade includes an airfoil comprising an open-cell solid ceramic foam comprising ceramic cell walls, and an intracellular volume therebetween. All or a part of each blade may be the open-cell solid ceramic foam, with the remainder metal.
The disk preferably comprises a disk nickel-base superalloy. The blades comprise the ceramic cell walls, preferably of alumina, and the intracellular volume, which may be partially a blade metal such as a blade nickel-base superalloy and partially empty porosity to reduce weight. The ceramic foam is preferably at least about 60 percent by volume of the ceramic cell walls, with the balance the intracellular volume.
The disk is preferably formed by conventional disk-fabrication technology. The blades are preferably formed by providing a piece of a sacrificial ceramic having the shape of the blade, and contacting the piece of the sacrificial ceramic to a reactive metal which reacts with the sacrificial ceramic to form an oxidized ceramic compound of the reactive metal and a reduced form of the ceramic. The resulting structure comprises the ceramic foam of the oxidized ceramic compound of the reactive metal with ceramic cell walls and the intracellular volume between the ceramic cell walls having a metallic reaction product therein. In subsequent processing, a portion of the metallic reaction product may be removed and/or replaced with another metal that is more suitable for the turbine blade application. The blades are conveniently joined to the disk by interdiffusion of the disk metal and the blade metal, but other techniques such as welding may be used as well.
The resulting structure has blades of the ceramic foam joined to the disk, which is preferably made of the disk nickel-base superalloy. The blades may have the entire intracellular volume filled with an appropriate metal such as the blade nickel-base superalloy. The blade nickel-base superalloy is integrally joined to the disk nickel-base superalloy by interdiffusion, saving the weight of a mechanical joint. The intracellular volume toward the tip of the turbine blade may be removed to leave porosity. The porosity reduces the weight of the turbine blade, and the resulting porous structure is resistant to erosion and impact damage. Because this portion of the turbine blade has no metal, it need not be protected by a protective coating structure such as a thermal barrier coating.
The present approach thus provides an integral blisk structure suitable for the high-pressure and low-pressure stages of a gas turbine engine. This structure achieves improved performance with reduced weight. Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.